According to one embodiment illustrated in FIG. 1, a fuselage frame 10 takes the form of a Z-section profile, of which the central portion referred to as the web 12 forms a complete or partial annulus. The profile comprises a first flange 14 known as the inner flange, positioned at the inner edge of the web 12 and perpendicular thereto, and a second flange 16 known as the outer flange positioned at the outer edge of the web 12 and also perpendicular thereto.
One method for producing such a frame in composite material is described in document FR-2.928.295.
According to that document, first of all a substantially rectangular strip is produced from a stack of three plies of pre-impregnated fibres, each ply having fibres oriented in one direction, the strip comprising plies with different fibre orientations, one ply 18 with fibres at 30°, one ply 20 with fibres at 90° and another ply 22 with fibres at 150°.
In a second stage, the strip of fibre plies is placed on a mandrel 24 made of deformable material and is then compressed on this mandrel so that it conforms to the shape thereof.
The deformable mandrel can be deformed between a rectilinear position and a curved position but has a cross section that is incompressible or near-incompressible.
Next, the deformed strip laid on the mandrel made of deformable material is brought into contact with a heated tooling fixture which at its periphery has radial sections with a profile that complements the cross sections of the mandrel. Thus, during bending, the strip is compressed and experiences an increase in temperature.
After this first strip has been placed, a second strip of thee plies of pre-impregnated fibres is cut out and placed on another deformable mandrel then compressed thereon.
Next, this second strip deformed on its mandrel made of deformable material is brought into contact with the first strip still in position on the tooling fixture and then compressed against the first strip.
To obtain a frame, it is necessary for several strips to be attached to one another as above, before the whole assembly thus formed is polymerized. To supplement this, plies with fibres oriented at 0° may be laid by hand in-between certain strips.
In the remainder of the description, a preform means a volume of pre-impregnated or not pre-impregnated fibres, derived notably from the stack of plies, layers, strips of fibres, which has not yet been polymerized.
This method of producing an aircraft frame may prove problematic during the stacking of the strips on one another.
It is problematic because the preform has a re-entrant angle, that surface of the preform that is not in contact with the tooling fixture having a concave shape. The same problem also arises when stacking plies in order to obtain, in composite, a rectilinear profile that has a re-entrant angle.
What happens is that when a layer is applied to a stack of layers, in order for the surfaces to be in contact over the entire width of the profile, the first point of contact needs to be situated at the vertex of the re-entrant angle, with a first side of the layer to be added to the stack being applied, by unrolling it, to the corresponding side of the stack that has been formed, with the same approach then being repeated on the second side.
Given the tack of the layers of pre-impregnated fibres, as soon as the layers come into contact with one another it is very difficult to separate them in order to readjust their positions. As a result, obtaining a preform with layers that are in contact across the entire width of the preform is a very tricky business. This difficulty is further compounded when the layers are not rectilinear but curved, as they are in the case of an aircraft frame.
Given these placement difficulties, a defect referred to as bridging may occur at the re-entrant angle, the two layers 26, 26′ not being pressed firmly against one another but delimiting a space 28 across the re-entrant angle, as illustrated in FIG. 5A.
When all the layers have been stacked, a preform 30 of pre-impregnated fibres is obtained which needs to undergo a polymerization cycle in order to obtain a composite component. As illustrated in FIG. 3, during this cycle, the preform 30 is positioned in a tooling fixture comprising a mould 32 on which the preform is positioned, and a covering 34 which covers the preform.
According to one embodiment, the mould 32 comprises means for extracting the gases, which means open via at least one port 36 onto a placement surface 38 on which the preform is placed, outside of the zone covered by the preform, but a short distance away from the said zone.
The covering 34 comprises:                a forming tool 40,        breather fabrics 42 provided at the periphery of the preform 30 and of the forming tool 40, in contact with the placement surface 38 at the ports 36 of the gas extraction means,        a release film 44 covering the forming tool 40,        a bleeder felt 46 which covers the forming tool 40 and the breather fabrics 42, and        a bag 48 which is attached to the placement surface 38 via sealing means 50 around the periphery of the breather fabrics 42.        
After this covering has been fitted, the preform undergoes a polymerization cycle at the end of which the fibres have become embedded in a matrix of resin. As illustrated in FIG. 4, the polymerization cycle comprises a temperature cycle 52, a pressure cycle 54 outside the bag and a vacuum-pulling cycle 56 inside the bag 48.
The temperature cycle 52 comprises a temperature increase, a temperature soak and a cooling phase. According to the procedure illustrated in FIG. 4, the temperature cycle comprises two temperature soaks.
The pressure cycle 54 comprises a pressurizing, a pressure hold and finally, a reduction in pressure down to atmospheric pressure.
The vacuum-pulling cycle 56 is started at the same moment as the temperature and pressure cycles and is maintained until the end of the cooling phase.
As pressurization is markedly more rapid than the increase in temperature, the bag 48 transmits the maximum pressure to the forming tool 40 for most of the temperature-increase time.
As illustrated in FIG. 5A, the forming tool 40 applies compressive load 58 to the first layer and this is then transmitted successively to the layers below. In the event of a defect, as illustrated in FIG. 5A, these compressive loads 58 which are almost normal to the contact surface create inter-layer shear stresses 60, notably between the layers 26, 26′ that are separated by the space 28.
When the temperature increases, adhesion between the layers decreases such that, because of the inter-layer shear stresses 60, there is tearing between two successive layers, leading to the formation of wrinkles 62 in the fibres as has been illustrated in FIG. 5B.
Thus, according to the prior art, the bridging type defect when stacking the layers leads to a defect whereby there are undulations in the fibres of the component at the end of the manufacturing process.
The presence of these undulations in the fibres in the composite component has a tendency to reduce the mechanical properties thereof.